The present invention relates to the field of turbomachines such as a turbojet engine or a turboprop engine of an aeroplane, and more particularly to a sealing device for a turbine of such a turbomachine.
Such a turbomachine having a front fan and two spools, for example, comprises, from upstream to downstream, a fan, a low-pressure compressor, a high-pressure compressor, a combustion chamber, a high-pressure turbine and a low-pressure turbine.
By convention, in the present application the terms “upstream” and “downstream” are defined in relation to the flow direction of the air in the turbojet engine. Similarly, by convention, in the present application the terms “inside” and “outside”, “lower” and “upper” and “inner” and “outer” are defined radially in relation to the axis of the engine. Thus, a cylinder extending along the axis of the engine comprises an inside face turned towards the axis of the engine and an outside surface, opposed to its inside surface.
A turbomachine low-pressure turbine comprises a turbine shaft on which are mounted a plurality of successive stages, each comprising an impeller and a nozzle. Each impeller comprises a disc carrying at its outer periphery substantially radial blades, the discs of the different impellers being connected coaxially to one another and to the drive shaft of the rotor of the turbine by suitable means. Each nozzle comprises an inner annular platform and an outer annular platform, between which extend substantially radial vanes. The outer platform of the nozzle comprises means for hooking and fixing on an outer casing of the turbine. The assembly of the nozzles forms the stationary portion of the motor known as the stator.
The inner platform comprises a wall extending radially towards the inside of the turbine and an annular ring connected to the wall. The wall and the annular ring thus define an upstream cavity and a downstream cavity between the turbine nozzle and the turbine rotor. The ring, the radial wall and the inner platform of the nozzle are generally formed of a single casting.
In operation, the vanes of the nozzle are exposed to the hot gases flowing in the flow path of the turbine. The temperature of the gases in the flow path is relatively high, typically in the order of 900 degrees Celsius, whereas the temperature in the region between the inner platform of the nozzle and the rotor is lower and for example around 700 degrees Celsius.
In order to consolidate this difference in temperature between the flow path and the cavity, which amounts to limiting the recycling of air, a turbine rotor is known which comprises an upstream spoiler defining a sealing baffle in the upstream cavity defined by the wall and the annular ring so as to form a baffle-type seal.
In order to ensure sealing which makes it possible to avoid heating the discs of the rotor, which would entail a potential risk to the integrity of the machine, and drops in the performance of the turbine such as flow-rate losses or generation of vortices, the downstream spoiler of the turbine rotor and the upstream portion of the inner platform, for each stage, overlap axially in part. Good sealing is thus ensured when the clearance between the downstream spoiler of the turbine rotor and the upstream portion of the inner platform, known as overlap clearance, is less than zero during operation of the turbine.
Another problem arises when the turbine overspeeds. The overspeed problem arises when considering, for example, breakage of the low-pressure turbine shaft during operation. This brings about a large increase in the rotational speed of the turbine, known as overspeed. To stop this malfunction, which can have serious consequences for the integrity of the turbine and thus of the aircraft, it is desirable to break a maximum number of runner blades on the nozzle blades on the basis that the rotor assembly will slip in relation to the stator (comprising the assembly of the nozzles) under the effect of the aerodynamic forces, bringing about contact between the stationary blades of nozzles of the stator and the moving blades of the rotor.
A distinction is made between harmful overspeed clearances, located between the trailing edge of the moving blades of the rotor and the leading edge of the nozzle at the upper level of or in the middle of the flow path, and non-harmful overspeed clearances, located between the trailing edge of the moving blades of the rotor and the leading edge of the nozzle at the lower level of the flow path.
To best control the overspeed and thus maintain the integrity of the turbine, the harmful overspeed clearances must first be bridged, that is to say become zero, before the non-harmful overspeed clearances. This means that the non-harmful overspeed clearances must be increased in order to ensure this hierarchy.
As shown in FIG. 1 and described below with reference to said FIG. 1, the following contradiction is encountered: an increase in the non-harmful lower flow path overspeed clearances involves a reduction in the overlap clearance and vice versa.